The present invention relates generally to gas turbine engines, and, more specifically, to turbine rotor blade vibration.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the gases in a high pressure turbine (HPT) which powers the compressor through one drive shaft, and additional energy is extracted from the gases in a low pressure turbine (LPT) which follows the HPT and drives another shaft for powering an upstream fan in a turbofan aircraft engine application.
Each turbine stage includes a stationary turbine nozzle having a row of nozzle vanes which guide the combustion gases through a corresponding row of turbine rotor blades extending radially outwardly from a supporting rotor disk. The rotor blades extract energy from the gases to rotate their supporting disks and the corresponding drive shaft connected thereto.
Each turbine blade is therefore subject to substantial aerodynamic pressure loads from the combustion gases, thermal loads from the heat thereof, and centrifugal loads from rotation of the blades atop their supporting rotating disks. The turbine blades are typically hollow in the initial turbine stages and include corresponding internal cooling circuits through which air bled from the compressor is channeled for cooling the blades from exposure to the hot combustion gases.
Each turbine rotor blade is therefore highly loaded during operation and is subject to vibration therefrom. Vibration occurs at distinct modes subject to excitation frequency and force and affects the high cycle fatigue (HCF) life of the rotor blades.
Accordingly, turbine rotor blades are specifically designed for their specific turbine stages and specific operating environment to minimize vibration at the different modes of vibration, and correspondingly maximize the HCF life of the blade.
In many designs, the individual rotor blades may be sufficiently configured for acceptable HCF life without additional remedy.
In other designs, a discrete damper is used for frictionally damping vibration of the blades during operation for enhancing blade life. However, dampers are generally undesirable because they increase the number of parts required for the engine, increase weight, and increase original cost of the engine as well as maintenance costs.
Blade vibration dampers are found in various configurations including those specifically configured for being mounted external to the blade, and those specifically configured for being mounted inside the blade. The different designs require different configurations and have different advantages and disadvantages and different modes of operation except for the common use of frictional damping.
A frictional damper introduces an interface centrifugally loaded during operation for effecting frictional damping as the adjacent components experience relative motion during vibration. The energy of vibration is dissipated by the friction, which therefore reduces the magnitude of the vibration.
However, frictional damping occurs with frictional wear between the components, and the damper and the associated blade being dampened must also be suitably designed for minimizing friction wear to ensure the desired useful life of the blade and cooperating damper.
Development testing of certain turbine rotor blades indicates that under-platform dampers are not suitably effective for damping vibration under certain vibratory modes. However, internal damping may be used to more effectively dampen the experienced vibratory modes, but substantially increases the difficulty of design.
Since the typical turbine rotor blade is optimized in design for aerodynamic, thermodynamic, and mechanical performance, the redesign thereof for additional damping performance necessarily affects the original optimum design.
For example, the airfoil portion of the typical turbine blade is hollow with relatively thin sidewalls, and includes an intricate internal cooling circuit differently configured for the different heat loads experienced over the opposite pressure and suction sides thereof. The experienced heat loads vary from the airfoil leading edge which first receives the hot combustion gases to the relatively thin trailing edge over which the gases are discharged.
The airfoil pressure side is generally concave and the suction side is generally convex and effect different velocity and pressure distributions thereover between the leading and trailing edges of the airfoil and from root to tip.
The introduction of an additional damper inside the airfoil therefore affects the cooling performance of the internal cooling circuit, as well as increases the weight of the blade and the corresponding centrifugal loads and stresses generated during rotary operation of the blades atop the supporting rotor disk.
Accordingly, it is desired to provide a turbine rotor blade having an internal damper for reducing blade vibration during operation while minimizing adverse affect in the overall blade design.